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A paper describes algorithms for guidance and control (G&C) of a spacecraft maneuvering near a planet, moon, asteroid, comet, or other small astronomical body. The algorithms were developed following a model-predictive-control approach along with a convexification of the governing dynamical equations, control constraints, and trajectory and state constraints.
During the Orion landing and recovery subsystem design review, June 2009, it was noted that the human system and various vehicle systems, the environmental control and life support (ECLSS) and guidance, navigation and control (GN&C) systems for example, are negatively affected by Orion assuming a stable 2 (upside down; Figure A) configuration post landing. The stable 2 configuration is predicted to occur about 50% of the time based on Apollo landing data and modeling of the current capsule. The stable 2 configuration will be countered by an active up-righting system (crew module up-righting system; CMUS). Post landing balloons will deploy and inflate causing the vehicle to assume or maintain the stable 1 (up-right; Figure B) configuration. During the design review it was proposed that the up-righting system could be capable of righting the vehicle within 60 seconds. However, this time limit posed a series of constraints on the design which made it less robust than desired. The landing and recovery subsystem team requested an analysis of Orion vehicle systems as well as the human system with regard to the effect of stable 2 in order to determine if an up-righting response time greater than 60 seconds could be tolerated. The following report focuses on the assessment of the human system in the posture assumed when Orion is in the stable 2 configuration. Stable 2 will place suited, seated, and restrained crewmembers in a prone (facedown), head-up position for a period of time dependent on the functionality of the up-righting systems, ability of the crew to release themselves from the seat and restraints, and/or time to arrival of rescue forces. Given that the Orion seat and restraint system design is not complete and therefore, not available for evaluation, Space Medicine assessed how long a healthy but deconditioned crewmember could stay in this prone, restrained position and the physiological consequences of this posture by researching terrestrial analogs and considered the known physiological alterations and deconditioning experienced by long duration crewmembers.
The primal-dual interior-point algorithm implemented in G-OPT is a relatively new and efficient way of solving convex optimization problems. Given a prescribed level of accuracy, the convergence to the optimal solution is guaranteed in a predetermined, finite number of iterations. G-OPT Version 1.0 is a flight software implementation written in C. Onboard application of the software enables autonomous, real-time guidance and control that explicitly incorporates mission constraints such as control authority (e.g. maximum thrust limits), hazard avoidance, and fuel limitations. This software can be used in planetary landing missions (Mars pinpoint landing and lunar landing), as well as in proximity operations around small celestial bodies (moons, asteroids, and comets). It also can be used in any spacecraft mission for thrust allocation in six-degrees-of-freedom control.
In the 1970 s the role of the military helicopter evolved to encompass more demanding missions including low-level nap-of-the-earth flight and operation in severely degraded visual environments. The Vertical Motion Simulator (VMS) at the NASA Ames Research Center was built to provide a high-fidelity simulation capability to research new rotorcraft concepts and technologies that could satisfy these mission requirements. The VMS combines a high-fidelity large amplitude motion system with an adaptable simulation environment including interchangeable and configurable cockpits. In almost 30 years of operation, rotorcraft research on the VMS has contributed significantly to the knowledge-base on rotorcraft performance, handling qualities, flight control, and guidance and displays. These contributions have directly benefited current rotorcraft programs and flight safety. The high fidelity motion system in the VMS was also used to research simulation fidelity. This research provided a fundamental understanding of pilot cueing modalities and their effect on simulation fidelity.
The Advanced Concepts Office at Marshall Space Flight Center completed a brief spacecraft design study for the 8-meter monolithic Advanced Technology Large Aperture Space Telescope (ATLAST-8m). This spacecraft concept provides all power, communication, telemetry, avionics, guidance and control, and thermal control for the observatory, and inserts the observatory into a halo orbit about the second Sun-Earth Lagrange point. The multidisciplinary design team created a simple spacecraft design that enables component and science instrument servicing, employs articulating solar panels for help with momentum management, and provides precise pointing control while at the same time fast slewing for the observatory.
On October 28th, 2009, the National Aeronautics and Space Administration (NASA) launched the Ares I-X Flight Test Vehicle (FTV) from pad 39B, providing the first set of flight test data for NASA's Ares I vehicle design team. This test was critical in providing insight into areas were significant design challenges existed. This paper discusses the objectives of the mission and how they were satisfied. It discusses the overall results of the flight test and look at the data retrieved from the flight. Ares I-X was highly instrumented with over 700 channels of Developmental Flight Instrumentation (DFI). Significant insight was gained in the areas of thrust oscillation, vibro-acoustics, predicting jet interactions and slag ejection from solid rocket systems with submerged nozzles. The paper outlines results from the Guidance Navigation & Control (GN&C), Thermal, Vibro-acoustic, Structures, Aero, Aero-Acoustic and Trajectory teams.
Current plans for the space station call for the GPS receiver to be installed on the U.S. lab module of the station in early 2001 (ISS Assembly Flight SA), followed by the attachment of the antenna array in late 2001 (Flight 8A). At that point the U.S. ISS guidance and control system will be operational. The flight of SIGI on the space station represents a "coming of age" for GPS technology on spacecraft. For at least a decade, the promise of using GPS receivers to automate spacecraft operations, simplify satellite design, and reduce mission costs has enticed satellite designers. Integration of this technology onto spacecraft has been slower than some originally anticipated. However, given the complexity of the GPS sensor, and the importance of the functions it performs, its incorporation into mainstream satellite design has probably occurred at a very reasonable pace. Going from providing experimental payloads on small, unmanned satellites to performing critical operational functions on manned vehicles has been a major evolution. If all goes as planned in the next few months, GPS receivers will soon provide those critical functions on one of the most complex spacecraft in history, the International Space Station.
The original G-Guidance algorithm provided an autonomous guidance and control policy for small-body proximity operations that took into account uncertainty and dynamics disturbances. However, there was a lack of robustness in regards to object proximity while in autonomous mode. The modified GGuidance algorithm was augmented with a second operational mode that allows switching into a safety hover mode. This will cause a spacecraft to hover in place until a mission-planning algorithm can compute a safe new trajectory. No state or control constraints are violated. When a new, feasible state trajectory is calculated, the spacecraft will return to standard mode and maneuver toward the target. The main goal of this augmentation is to protect the spacecraft in the event that a landing surface or obstacle is closer or further than anticipated. The algorithm can be used for the mitigation of any unexpected trajectory or state changes that occur during standard mode operations
The Orion Multi-Purpose Crew Vehicle (MPCV) will perform a flight test known as Exploration Flight Test-1 (EFT-1) currently scheduled for 2014. One of the primary functions of this test is to exercise all of the important Guidance, Navigation, Control (GN&C), and Propulsion systems, along with the flight software for future flights. The Descent and Landing segment of the flight is governed by the requirements levied on the GN&C system by the Landing and Recovery System (LRS). The LRS is a complex system of parachutes and flight control modes that ensure that the Orion MPCV safely lands at its designated target in the Pacific Ocean. The Descent and Landing segment begins with the jettisoning of the Forward Bay Cover and concludes with sensing touchdown. This paper discusses the requirements, design, testing, analysis and performance of the current EFT-1 Descent and Landing Triggers flight software.
This paper studies the Touch-and-Go (TAG) concept for enabling a spacecraft to take a sample from the surface of a small primitive body, such as an asteroid or comet. The idea behind the TAG concept is to let the spacecraft descend to the surface, make contact with the surface for several seconds, and then ascend to a safe location. Sampling would be accomplished by an end-effector that is active during the few seconds of surface contact. The TAG event is one of the most critical events in a primitive body sample-return mission. The purpose of this study is to evaluate the dynamic behavior of a representative spacecraft during the TAG event, i.e., immediately prior, during, and after surface contact of the sampler. The study evaluates the sample-collection performance of the proposed sampling end-effector, in this case a brushwheel sampler, while acquiring material from the surface during the contact. A main result of the study is a guidance and control (G&C) validation of the overall TAG concept, in addition to specific contributions to demonstrating the effectiveness of using nonlinear clutch mechanisms in the sampling arm joints, and increasing the length of the sampling arms to improve robustness.
The Geostationary Operational Environmental Satellite-R Series (GOES-R) is the first of the next generation geostationary weather satellites, scheduled for delivery in late 2015. GOES-R represents a quantum increase in Earth and solar weather observation capabilities, with 4 times the resolution, 5 times the observation rate, and 3 times the number of spectral bands for Earth observations. With the improved resolution, comes the instrument suite's increased sensitive to disturbances over a broad spectrum 0-512 Hz. Sources of disturbance include reaction wheels, thruster firings for station keeping and momentum management, gimbal motion, and internal instrument disturbances. To minimize the impact of these disturbances, the baseline design includes an Earth Pointed Platform (EPP), a stiff optical bench to which the two nadir pointed instruments are collocated together with the Guidance Navigation & Control (GN&C) star trackers and Inertial Measurement Units (IMUs). The EPP is passively isolated from the spacecraft bus with Honeywell D-Strut isolators providing attenuation for frequencies above approximately 5 Hz in all six degrees-of-freedom. A change in Reaction Wheel Assembly (RWA) vendors occurred very late in the program. To reduce the risk of RWA disturbances impacting performance, a secondary passive isolation system manufactured by Moog CSA Engineering was incorporated under each of the six 160 Nms RWAs, tuned to provide attenuation at frequencies above approximately 50 Hz. Integrated wheel and isolator testing was performed on a Kistler table at NASA Goddard Space Flight Center. High fidelity simulations were conducted to evaluate jitter performance for four topologies: 1) hard mounted no isolation, 2) EPP isolation only, 2) RWA isolation only, and 4) dual isolation. Simulation results demonstrate excellent performance relative to the pointing stability requirements, with dual isolated Line of Sight (LOS) jitter less than 1 micron rad.
This paper describes the landmark measurement system being developed for terrain relative navigation on NASAs Asteroid Redirect Robotic Mission (ARRM),and the results of a performance characterization study given realistic navigational and model errors. The system is called Retina, and is derived from the stereo-photoclinometry methods widely used on other small-body missions. The system is simulated using synthetic imagery of the asteroid surface and discussion is given on various algorithmic design choices. Unlike other missions, ARRMs Retina is the first planned autonomous use of these methods during the close-proximity and descent phase of the mission.
As a part of the NASA initiative to collect a boulder off of an asteroid and return it to Lunar orbit, the Satellite Servicing Capabilities Office (SSCO) and NASA GSFC are developing an on-board relative terrain imaging navigation algorithm for the Asteroid Redirect Robotic Mission (ARRM). After performing several flybys and dry runs to verify and refine the shape, spin, and gravity models and obtain centimeter level imagery, the spacecraft will descend to the surface of the asteroid to capture a boulder and return it to Lunar Orbit. The algorithm implements Stereophotoclinometry methods to register landmarks with images taken onboard the spacecraft, and use these measurements to estimate the position and orientation of the spacecraft with respect to the asteroid. This paper will present an overview of the ARRM GN&C system and concept of operations as well as a description of the algorithm and its implementation. These techniques will be demonstrated for the descent to the surface of the proposed asteroid of interest, 2008 EV5, and preliminary results will be shown.
This paper describes the landmark measurement system being developed for terrain relative navigation on NASAs Asteroid Redirect Robotic Mission (ARRM),and the results of a performance characterization study given realistic navigational and model errors. The system is called Retina, and is derived from the stereophotoclinometry methods widely used on other small-body missions. The system is simulated using synthetic imagery of the asteroid surface and discussion is given on various algorithmic design choices. Unlike other missions, ARRMs Retina is the first planned autonomous use of these methods during the close-proximity and descent phase of the mission.
Flight testing and modeling techniques were developed to accurately identify global nonlinear aerodynamic models for aircraft in real time. The techniques were developed and demonstrated during flight testing of a remotely-piloted subscale propeller-driven fixed-wing aircraft using flight test maneuvers designed to simulate a Learn-To-Fly scenario. Prediction testing was used to evaluate the quality of the global models identified in real time. The real-time global nonlinear aerodynamic modeling algorithm will be integrated and further tested with learning adaptive control and guidance for NASA Learn-To-Fly concept flight demonstrations.
The Geostationary Operational Environmental Satellite-R Series (GOES-R) is the first of the next generation geostationary weather satellites. GOES-R successfully launched on November19, 2016 and renamed GOES-16 upon entering geostationary orbit. Subsequently, GOES-16post-launch testing began. This paper presents the GOES-16 Satellite Dynamic Interaction Characterization results for the Earth Pointed Platform (EPP) stowed, referred to as the Reaction Wheel Assembly (RWA) Isolation Only configuration, and deployed, referred to as the Dual Isolation configuration. GOES-R represents a quantum increase in Earth and solar weather observation capabilities, with 4 times the resolution, 5 times the observation rate, and 3 times the number of spectral bands for Earth observations. With the improved resolution, comes the instrument suites increased sensitivity to disturbances over a broad spectrum 0-512Hz. Sources of disturbance include reaction wheels, thruster firings for station keeping and momentum management, gimbal motion, and internal instrument disturbances. To minimize the impact of these disturbances, the baseline design included an EPP, a stiff optical bench to which the two nadir pointed instruments are collocated together with the Guidance Navigation Control (GNC) star trackers and Inertial Measurement Units (IMUs). The EPP is passively isolated from the spacecraft bus with Honeywell D-Strut isolators providing attenuation for frequencies above 5 Hz in all six degrees-of-freedom. To reduce the risk of wheel disturbances impacting performance, a secondary passive isolation system manufactured by Moog CSA Engineering was incorporated under each of the six 160 Nms reaction wheels, tuned to provide attenuation at frequencies above 50 Hz. Integrated wheel and isolator testing was performed on a Kistler table at NASA Goddard Space Flight Center. Pre-launch Satellite Dynamic Interaction Characterization high-fidelity simulations and ground testing were conducted to evaluate jitter performance for two cases: 1) deployed EPP and reaction wheel (Dual Isolation) and 2) EPP hard mounted (RWA Isolation Only) to the spacecraft. A comparison of pre-launch to post-launch Satellite Dynamic Interaction Characterization results are also presented in this paper.
NASA is developing the Near Earth Asteroid (NEA) Scout mission that will use a solar sail to travel to an asteroid where it will perform a slow flyby to acquire science imagery. A guidance and control system was developed to meet the science and trajectory requirements. The NEA Scout design process can be applied to an interstellar or precursor mission that uses a beam-propelled sail. The scientific objectives are met by accurately targeting the destination trajectory position and velocity. The destination is targeted by understanding the force on the sail from the beam (or sunlight in the case of NEA Scout) over the duration of the thrust maneuver. The propulsive maneuver is maintained by accurate understanding of the torque on the sail, which is a function of sail shape, optical properties, and mass properties, all of which apply to NEA Scout and beam propelled sails. NEA Scout uses active control of the sail attitude while trimming the solar torque, which could be used on a beamed propulsion sail if necessary. The biggest difference is that NEA Scout can correct for uncertainties in sail thrust modeling, spacecraft orbit, and target orbit throughout the flight to the target, while beamed propulsion needs accurate operation for the short duration of the beamed propulsion maneuver, making accurate understanding of the sail thrust and orbits much more critical.
The Mid-Lift-to-Drag Ratio Rigid Vehicle (MRV) is a proposed candidate in the NASA Evolvable Mars Campaign's (EMC) Pathfinder Entry, Descent, and Landing (EDL) architecture study. The purpose of the study is to design a mission and vehicle capable of transporting a 20mt payload to the surface of Mars. The MRV is unique in its rigid, asymmetrical lifting-body shape which enables a higher lift-to-drag ratio (L/D) than the typical robotic Mars entry capsule vehicles that carry much less mass. This paper presents the formulation and six-degree-of-freedom (6DOF) performance of the MRV's control system, which uses both aerosurfaces and a propulsive reaction control system (RCS) to affect longitudinal and lateral directional behavior.